A gas turbine engine includes a compression section, a combustor, and a turbine. Air flows axially through the engine. As is well known in the art, the air is compressed in the compressor, where its temperature and pressure are raised. After being discharged from the compressor, the air enters the combustor, is mixed with fuel and burned therein. The hot products of combustion emerging from the combustor enter the turbine, where the hot gases expand to produce engine thrust and to drive the turbine, which in turn drives the compressor.
Both the compression and turbine sections include alternating rows or stages of rotating and stationary airfoils. Each airfoil comprises an airfoil portion having a leading edge and a trailing edge. The components of the turbine operate in an especially hostile environment that is characterized by extremely high temperatures. The temperature of hot combustion gases entering the turbine generally exceeds the melting point temperatures of the alloys from which the turbine airfoils are fabricated. Thus, to properly perform in such a harsh environment, the turbine airfoils must be cooled. The initial stages of turbine airfoils need substantially more cooling than subsequent stages thereof because the temperature and pressure of gaseous products of combustion are highest at the turbine entrance, and decrease progressively therefrom. Moreover, each airfoil requires more cooling at the leading edge than at the trailing edge because the temperature and pressure of the products of combustion are higher at the leading edge of the airfoil than at the trailing edge thereof.
Generally, turbine airfoil cooling is accomplished by internal impingement cooling, internal convection cooling, or some combination thereof. In convection cooling, cooling air flows through a typically serpentine passage within the airfoil, continuously removing heat therefrom. With impingement cooling, cooling air is channeled to the inside of the airfoil and directed against the inside walls of the airfoil. The air then exits the airfoil through a set of film holes provided within the airfoil walls. Although both convection cooling and air impingement cooling are effective methods for cooling blades, the impingement cooling has a higher pressure drop associated therewith.
Frequently, multiple cooling passages supplied from a single source of cooling air are used in a single airfoil. A second stage turbine stationary airfoil (vane), comprising an airfoil portion flanged by an inner diameter platform and an outer diameter platform, is convectionally cooled with a multi-pass serpentine type passage within the trailing edge and a single-pass passage within the leading edge. The air from the serpentine passage is discharged through a plurality of slots disposed near the trailing edge, whereas the air from the leading edge passage is discharged through an outlet disposed in the inner platform and is subsequently utilized to cool an inner seal which is disposed radially inward of the inner platform of the second stage stationary airfoil. The air discharged from the leading edge passage is still cool relative to the temperature of the inner seal and provides adequate cooling thereto.
The cooling air is bled from the engine's compressor and bypasses the combustor. It will be understood that any compressor bleed air diverted from the compressor for such cooling will be unavailable to support combustion in the combustor and therefore will reduce engine power. Thus, to minimize any sacrifice in engine performance due to reduced airflow to support combustion, any scheme for cooling turbine airfoils must optimize the use of compressor bleed cooling air. Furthermore, it is desirable to bleed compressor air for cooling from the initial stages of the compressor rather than from the final stages thereof for two reasons. First, at the initial stages of the compressor, less work has been invested in compressing air than at the final stages of the compressor and thus, engine power losses due to cooling requirements are lesser. Secondly, compressor air at initial stages has a lower temperature than at the later stages, which is desirable for cooling. However, air from the initial stages of the compressor may not have sufficient pressure to provide adequate cooling for the turbine airfoils. Thus, there is a trade-off between either bleeding lower temperature, lower pressure compressor air from initial stages without having invested too much work into compressing the air, but not getting sufficient pressure for cooling; or bleeding higher temperature, higher pressure compressor air from final stages with reduced engine performance.
It has been determined that the above mentioned cooling scheme for the second stage turbine vane must be improved upon for modem, higher thrust engines. Modem compressors include additional stages of airfoils resulting in higher temperatures and pressures of the products of combustion entering the turbine. The turbine airfoils, in general, need more cooling than prior art airfoils to avoid burning of the walls thereof. In particular, the leading edge of the airfoils needs more cooling than the trailing edge thereof because the external pressure and temperatures are higher at the leading edge. Impingement cooling would provide a better cooling for the leading edge than convection cooling. The drawback to impingement cooling of the leading edge is that the pressure drop of the cooling air in the passage is higher than when convection cooling is used. A high pressure drop results in lower internal pressure in the leading edge than in the trailing edge. Additionally, the external pressure at the leading edge is higher than the external pressure at the trailing edge. If the pressure of the airfoil internal air is lower than the pressure of external air (products of combustion), the hot external air will enter the internal passage and burn the walls of the airfoil. Thus, to avoid inward flow of external air, the leading edge requires higher internal pressure than the trailing edge.
Although cooling by higher pressure compressor air is necessary for the airfoil to avoid burning as set forth hereinabove, there are two major drawbacks to the use of higher pressure compressor air. Use of higher pressure compressor air exclusively would be undesirable because it directly diverts air from the combustion process, thereby reducing the overall performance of the engine. Secondly, the compressor bleed air with higher pressure also has higher temperature. The temperature of the higher pressure compressor air would be too high to provide adequate cooling for the inner seal. Thus, there is an acute need to provide sufficient cooling to modem turbine airfoils without incurring a penalty in engine performance.